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GOE 416A AIRFOIL (goe416a-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 416A AIRFOIL (goe416a-il)
Reynolds number: 200,000
Max Cl/Cd: 61.35 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe416a-il-200000.txt
Download as CSV file: xf-goe416a-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 416A AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.750  -0.4555   0.13348   0.13010  -0.0026   1.0000   0.0437
 -12.500  -0.4580   0.12984   0.12648  -0.0043   1.0000   0.0458
 -12.250  -0.4666   0.12526   0.12192  -0.0074   1.0000   0.0477
  -6.000  -0.5366   0.02447   0.01724  -0.0334   1.0000   0.0446
  -5.750  -0.5103   0.02359   0.01602  -0.0314   1.0000   0.0365
  -5.500  -0.4883   0.02174   0.01410  -0.0301   1.0000   0.0362
  -5.250  -0.4665   0.02047   0.01275  -0.0288   1.0000   0.0364
  -5.000  -0.4454   0.01888   0.01117  -0.0274   1.0000   0.0354
  -4.750  -0.4246   0.01768   0.00997  -0.0261   1.0000   0.0348
  -4.500  -0.4038   0.01668   0.00897  -0.0249   1.0000   0.0347
  -4.250  -0.3823   0.01581   0.00808  -0.0240   1.0000   0.0351
  -4.000  -0.3599   0.01506   0.00728  -0.0234   1.0000   0.0359
  -3.750  -0.3365   0.01441   0.00652  -0.0231   1.0000   0.0383
  -3.500  -0.3122   0.01381   0.00579  -0.0229   1.0000   0.0422
  -3.250  -0.2741   0.01337   0.00521  -0.0252   0.9963   0.0483
  -3.000  -0.2345   0.01118   0.00479  -0.0294   0.9922   0.4594
  -2.750  -0.1909   0.01182   0.00544  -0.0322   0.9843   0.5238
  -2.500  -0.1505   0.01219   0.00581  -0.0344   0.9759   0.5513
  -2.250  -0.1073   0.01242   0.00612  -0.0371   0.9701   0.5720
  -2.000  -0.0690   0.01260   0.00631  -0.0389   0.9609   0.5919
  -1.750  -0.0286   0.01249   0.00622  -0.0412   0.9531   0.6004
  -1.500   0.0089   0.01236   0.00600  -0.0431   0.9422   0.6072
  -1.250   0.0421   0.01214   0.00576  -0.0439   0.9286   0.6108
  -1.000   0.0727   0.01196   0.00556  -0.0440   0.9134   0.6142
  -0.750   0.1011   0.01182   0.00535  -0.0437   0.8969   0.6180
  -0.500   0.1277   0.01172   0.00515  -0.0432   0.8773   0.6218
  -0.250   0.1537   0.01156   0.00497  -0.0424   0.8577   0.6245
   0.000   0.1798   0.01145   0.00481  -0.0416   0.8381   0.6278
   0.250   0.2063   0.01138   0.00468  -0.0411   0.8166   0.6316
   0.500   0.2333   0.01134   0.00454  -0.0406   0.7968   0.6353
   0.750   0.2602   0.01127   0.00443  -0.0402   0.7752   0.6384
   1.000   0.2870   0.01124   0.00435  -0.0398   0.7542   0.6420
   1.250   0.3142   0.01126   0.00431  -0.0395   0.7327   0.6461
   1.500   0.3416   0.01131   0.00425  -0.0392   0.7113   0.6502
   1.750   0.3686   0.01133   0.00423  -0.0389   0.6890   0.6537
   2.000   0.3956   0.01139   0.00425  -0.0386   0.6681   0.6581
   2.250   0.4231   0.01150   0.00431  -0.0385   0.6486   0.6632
   2.500   0.4506   0.01157   0.00438  -0.0384   0.6294   0.6678
   2.750   0.4779   0.01166   0.00448  -0.0383   0.6114   0.6728
   3.000   0.5055   0.01179   0.00459  -0.0382   0.5939   0.6786
   3.250   0.5327   0.01185   0.00470  -0.0381   0.5752   0.6843
   3.500   0.5600   0.01192   0.00482  -0.0379   0.5554   0.6917
   3.750   0.5867   0.01199   0.00491  -0.0377   0.5354   0.6990
   4.000   0.6139   0.01206   0.00506  -0.0376   0.5139   0.7081
   4.250   0.6403   0.01214   0.00520  -0.0372   0.4934   0.7177
   4.500   0.6664   0.01225   0.00538  -0.0369   0.4703   0.7300
   4.750   0.6916   0.01239   0.00558  -0.0364   0.4430   0.7468
   5.000   0.7155   0.01252   0.00580  -0.0356   0.4128   0.7728
   5.250   0.7374   0.01252   0.00603  -0.0342   0.3841   0.8357
   5.500   0.7671   0.01257   0.00622  -0.0346   0.3505   1.0000
   5.750   0.7939   0.01294   0.00652  -0.0347   0.3176   1.0000
   6.000   0.8196   0.01340   0.00686  -0.0345   0.2838   1.0000
   6.250   0.8438   0.01403   0.00732  -0.0343   0.2513   1.0000
   6.500   0.8674   0.01473   0.00789  -0.0339   0.2158   1.0000
   6.750   0.8908   0.01544   0.00844  -0.0336   0.1620   1.0000
   7.000   0.9101   0.01679   0.00937  -0.0328   0.1032   1.0000
   7.250   0.9300   0.01807   0.01051  -0.0319   0.0755   1.0000
   7.500   0.9491   0.01941   0.01177  -0.0309   0.0570   1.0000
   7.750   0.9659   0.02104   0.01340  -0.0295   0.0464   1.0000
   8.000   0.9849   0.02232   0.01472  -0.0284   0.0397   1.0000
   8.250   1.0018   0.02403   0.01649  -0.0270   0.0353   1.0000
   8.500   1.0213   0.02543   0.01802  -0.0258   0.0323   1.0000
   8.750   1.0399   0.02702   0.01964  -0.0247   0.0301   1.0000
   9.000   1.0565   0.03026   0.02294  -0.0235   0.0281   1.0000
   9.250   1.0757   0.03158   0.02451  -0.0223   0.0267   1.0000
   9.500   1.0934   0.03331   0.02648  -0.0211   0.0251   1.0000
   9.750   1.1092   0.03568   0.02912  -0.0198   0.0243   1.0000
  10.000   1.1221   0.03832   0.03205  -0.0183   0.0236   1.0000
  10.250   1.1313   0.04129   0.03534  -0.0165   0.0233   1.0000
  10.500   1.1348   0.04467   0.03909  -0.0144   0.0232   1.0000
  10.750   1.1307   0.04846   0.04328  -0.0120   0.0233   1.0000
  11.000   1.1174   0.05224   0.04742  -0.0090   0.0236   1.0000
  11.250   1.0998   0.05625   0.05175  -0.0068   0.0239   1.0000
  11.500   1.0801   0.06073   0.05649  -0.0060   0.0242   1.0000
  11.750   1.0588   0.06582   0.06183  -0.0069   0.0245   1.0000
  12.000   1.0364   0.07171   0.06795  -0.0093   0.0249   1.0000
  12.250   1.0130   0.07854   0.07498  -0.0133   0.0252   1.0000
  12.500   0.9887   0.08634   0.08295  -0.0185   0.0255   1.0000
  12.750   0.9629   0.09540   0.09216  -0.0250   0.0258   1.0000
  13.000   0.9356   0.10584   0.10271  -0.0324   0.0262   1.0000
  13.250   0.9071   0.11774   0.11468  -0.0401   0.0268   1.0000
  13.500   0.8861   0.12794   0.12487  -0.0451   0.0277   1.0000
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